Method and assembly for reducing secondary heat in a gas turbine engine

ABSTRACT

A turbine section for a gas turbine engine includes a first rotor assembly with a first rotor assembly bleed air source and an aft cavity that is in fluid communication with the first rotor assembly bleed air source. A second rotor assembly includes a forward cavity. A vane bleed air source is in fluid communication with the forward cavity. A seal extends between the first rotor assembly and the second rotor assembly.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application claims priority to U.S. Provisional Application No.62/049,422, which was filed on Sep. 12, 2014 and is incorporated hereinby reference.

BACKGROUND

Gas turbine engines with multiple turbine stages include interstage sealarrangements between adjacent stages for improved operating efficiency.The interstage seal arrangements confine the flow of hot combustion coregases within an annular path around and between stationary turbinestator blades, nozzles, and also around and between adjacent rotorblades.

The interstage seal arrangements may also serve to confine and directcooling air to cool the turbine disks, the turbine blade roots, and alsothe interior of the rotor blades themselves as rotor blade coolingfacilities higher turbine inlet temperatures, which results in higherthermal efficiency of the engine and higher thrust output. Theinterstage seal configurations must also accommodate axial and radialmovements of the turbine stage elements during engine operation as theseveral elements are subjected to a range of different loadings anddifferent rates of expansion based upon local part temperatures andaircraft operating conditions.

One such interstage seal includes a monolithic interstage seal structurethat spans the axial gap between the rotor disks. Although effective,the monolithic interstage seal is typically manufactured from highstrength materials to withstand the relatively high hoop stressesinduced by rotation. This may result in a relatively heavy sealstructure that imposes additional weight adjacent the rotor disks.Another interstage seal is a segmented seal assembly in which multiplesegments are assembled together circumferentially. Although effective,the multiple segments may increase transient load variation between therotor disks and may result in additional leakage paths between thesegments.

SUMMARY

In one exemplary embodiment, a turbine section for a gas turbine engineincludes a first rotor assembly with a first rotor assembly bleed airsource and an aft cavity that is in fluid communication with the firstrotor assembly bleed air source. A second rotor assembly includes aforward cavity. A vane bleed air source is in fluid communication withthe forward cavity. A seal extends between the first rotor assembly andthe second rotor assembly.

In a further embodiment of the above, the first rotor assembly bleed airsource is at a first pressure and the vane bleed air source is at asecond pressure lower than the first pressure.

In a further embodiment of any of the above, the seal is a segmentedseal and includes a plurality of knife edges that mate with an abradablematerial on a radially inner end of a vane separating the first rotorassembly from the second rotor assembly.

In a further embodiment of any of the above, a radially inner edge ofthe abradable material is located between approximately 10% and 50% of aradial length from a radial inner edge of a root portion of a blade ofthe first rotor assembly.

In a further embodiment of any of the above, a radially inner end of thevane includes a radially extending bleed air passage.

In a further embodiment of any of the above, the first stage aft cavityincludes a radially inner edge that is located between approximately 10%and 50% of a radial length from a radial inner edge of a root portion ofa blade in the first rotor assembly.

In a further embodiment of any of the above, the second stage forwardcavity includes a radially inner edge located between approximately 10%and 50% of a radial length from a radial inner edge of a root portion ofa blade in the second rotor assembly.

In a further embodiment of any of the above, the first rotor assemblyincludes a first web with a radially innermost protrusion locatedradially outward at between approximately 30% and 60% of a radial lengthfrom a radial inner edge of the first web.

In a further embodiment of any of the above, a second stage has a secondweb with a radially innermost protrusion located radially outward atbetween 30% and 60% of a radial length from a radial inner edge of thesecond web.

In another exemplary embodiment, a gas turbine engine includes a turbinesection including a first rotor assembly that includes an aft cavitythat is in fluid communication with a first rotor assembly bleed airsource for the first rotor assembly. A second rotor assembly includes aforward cavity. A vane bleed air source is in fluid communication withthe forward cavity. A seal extends between the first rotor assembly andthe second rotor assembly.

In a further embodiment of the above, the first rotor assembly bleed airsource is at a first pressure and the vane bleed air source is at asecond pressure lower than the first pressure.

In a further embodiment of any of the above, the seal is a segmentedseal and includes a plurality of knife edges that mate with an abradablematerial on a radially inner end of a vane separating the first rotorassembly from the second rotor assembly.

In a further embodiment of any of the above, a radially inner edge ofthe abradable material is located between approximately 10% and 50% of aradial length from a radial inner edge of a root portion of a blade ofthe first rotor assembly.

In a further embodiment of any of the above, a radially inner end of thevane includes a radially extending bleed air passage.

In a further embodiment of any of the above, the first stage aft cavityincludes a radially inner edge located between approximately 10% and 50%of a radial length from a radial inner edge of a root portion of a bladein the first rotor assembly.

In a further embodiment of any of the above, the second stage forwardcavity includes a radially inner edge located between approximately 10%and 50% of a radial length from a radial inner edge of a root portion ofa blade in the second rotor assembly.

In another exemplary embodiment, a method of operating a gas turbineengine includes directing bleed air from a first rotor assembly bleedair source in a first rotor assembly at a first pressure to purge afirst stage aft cavity and directing bleed air from a vane bleed airsource in a vane at a second pressure to purge a second stage forwardcavity. The first pressure is greater than the second pressure.

In a further embodiment of the above, a radially inner end of the vaneincludes a radially extending bleed air passage.

In a further embodiment of any of the above, the first stage aft cavityincludes a radially inner edge located between approximately 10% and 50%of a radial length from a radial inner edge of a root portion of a bladein the first rotor assembly.

In a further embodiment of any of the above, the second stage forwardcavity includes a radially inner edge located between approximately 10%and 50% of a radial length from a radial inner edge of a root portion ofa blade in a second rotor assembly.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic view of an example gas turbine engine.

FIG. 2 is an enlarged schematic cross-section of a high pressure turbinesection for the gas turbine engine of FIG. 1.

FIG. 3 is a perspective view of an example seal segment.

FIG. 4 is another perspective view of the example seal segment of FIG.3.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a second (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 57 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axisA which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present disclosure isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft(10,668 meters), with the engine at its best fuel consumption—also knownas “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is theindustry standard parameter of lbm of fuel being burned divided by lbfof thrust the engine produces at that minimum point. “Low fan pressureratio” is the pressure ratio across the fan blade alone, without a FanExit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosedherein according to one non-limiting embodiment is less than about 1.45.“Low corrected fan tip speed” is the actual fan tip speed in ft/secdivided by an industry standard temperature correction of [(Tram °R)/(518.7° R)]^(0.5). The “Low corrected fan tip speed” as disclosedherein according to one non-limiting embodiment is less than about 1150ft/second (350.5 meters/second).

FIG. 2 illustrates an enlarged schematic view of the high pressureturbine 54, however, other sections of the gas turbine engine 20 couldbenefit from this disclosure. The high pressure turbine 54 includes atwo-stage turbine section with a first rotor assembly 60 and a secondrotor assembly 62. The first rotor assembly 60 and the second rotorassembly 62 are attached to the outer shaft 50.

The first rotor assembly 60 includes a first array of rotor blades 64circumferentially spaced around a first disk 68 and the second rotorassembly 62 includes a second array of rotor blades 66 circumferentiallyspaced around a second disk 70. Each of the first and second array ofrotor blades 64, 66 include a respective first root portion 72 and asecond root portion 74, a first platform 76 and a second platform 78,and a first airfoil 80 and a second airfoil 82. Each of the first andsecond root portions 72, 74 is received within a respective first rimand a second rim 84, 86 of the first and second disk 68, 70. The firstairfoil 80 and the second airfoil 82 extend radially outward toward afirst and second blade outer air seal (BOAS) assembly 81, 83,respectively.

The first and second array of rotor blades 64, 66 are disposed in thecore flow path that is pressurized in the compressor section 24 thenheated to a working temperature in the combustor section 26. The firstand second platforms 76, 78 separate a gas path side inclusive of thefirst and second airfoils 80, 82 and a non-gas path side inclusive ofthe first and second root portions 72, 74.

A shroud assembly 88 within the engine case structure 36 between thefirst rotor assembly 60 and the second rotor assembly 62 directs the hotgas core airflow in the core flow path from the first array of rotorblades 64 to the second array of rotor blades 66. The shroud assembly 88may at least partially support the first and second blade outer airseals 81, 83 and include an array of vanes 90 that extend between arespective inner vane platform 92 and an outer vane platform 94. Theouter vane platform 94 may be supported by the engine case structure 36and the inner vane platform 92 supports an abradable annular seal 96,such as a honeycomb, to seal the hot gas core airflow in the axialdirection with respect to a segmented interstage seal assembly 100.

The segmented interstage seal assembly 100 includes a plurality ofindividual seal segments 102 (FIG. 3) disposed between the first andsecond rotor assemblies 60, 62. The segmented interstage seal assembly100 creates a seal between the axially flowing hot gas core airflow anda radially inner volume 104 between a respective first and second web106, 108 and a first and second hub 110, 112 of the first and seconddisk 68, 70.

Each seal segment 102 is radially supported on a respective pilotdiameter 114, 116, formed by the respective first and second rim 84, 86of the first and second disk 68, 70. At least one of the individual sealsegments 102 includes an anti-rotation tab 118 that interfaces with astop 120 on the first rim 84 of the first disk 68. It should beappreciated that various interfaces may be alternatively or additionallyprovided on one, or multiple, seal segments 102. At least a portion ofthe seal segment extends continuously between the first rotor assembly60 and the second rotor assembly 62.

As shown in FIGS. 3 and 4, each seal segment 102 generally includes afirst circumferential end 130, a second circumferential end 132, a firstradial side 134, a second radial side 136, a first axial side 138, and asecond axial side 140. A plurality of knife edges 146 extend along thesecond radial side 136 for mating with the abradable annular seal 96(FIG. 2) on the inner vane platform 92 to reduce hot gas ingestionbetween the first and second rotor assemblies 60, 62.

A first plurality of openings 148 extend through the second radial side136 of the seal segment 102 and a second plurality of openings 150extend through a second axial side 140 of the seal segment to receiveand direct vane bleed air 152 a from a passage 152 through the vanes 90into the second disk 70. The second plurality of openings 150 therebyoperate essentially as a tangential onboard injector (TOBI).

A first stage aft cavity 154 is located downstream of the first rotorassembly 60 and is bounded by the first rotor assembly 60, the segmentedinterstage seal assembly 100 and the vane 90. A second stage forwardcavity 156 is located upstream of the second rotor assembly 62 and isbounded by the second rotor assembly 62, the segmented interstage sealassembly 100, and the vane 90.

During operation of the gas turbine engine 20, hot exhaust gases canenter the first stage aft cavity 154 and the second stage forward cavity156. The segmented interstage seal assembly 100 reduces the volume ofthe first stage aft cavity 154 and the second stage forward cavity 156and reduces the bleed air needed to purge and cool each of thesecavities. Therefore, the first stage aft cavity 154 can be purged bybleed air 157 from the first rotor assembly 60 at a first pressure andthe second stage forward cavity 156 can be purged by the vane bleed air152 a from the passage 152 in the vane 90. The first pressure of thebleed air 157 from the first rotor assembly 60 is greater than thesecond pressure of the bleed air from the vane 90.

The reduced volume of the first stage aft cavity 154 and the secondstage forward cavity 156 decreases secondary heat generation within thecavities 154 and 156 because there is less air that can be turned andheated during operation of the gas turbine engine 20. The first stageaft cavity 154 includes a radially inner edge 154 a located betweenapproximately 10% and 50% of the radial length from a radial inner edgeof the first root portion 72 of the first rotor assembly 60. The secondstage forward cavity 156 includes a radially inner edge 156 a locatedbetween approximately 10% and 50% of the radial length from a radialinner edge of the second root portion 74 of the second rotor assembly62. Additionally, a radially inner edge of the abradable annular seal 96is located between approximately 10% and 50% of a radial length from aradial inner edge of the first root portion 72 and between approximately10% and 50% of a radial length from a radial inner edge of the secondroot portion 74.

Secondary heat generation in the radially inner volume 104 is reduced bypositioning the protrusions, such as the stop 120, on radially outerportion of the first and second web 106, 108, thereby reducing turningof the air within the radially inner volume 104, reducing frictiongenerated heat associated with the turning. In the illustrated example,the protrusions are located radially outward at between approximately30% and 60% of a radial length from a radial inner edge of the first andsecond web 106, 108. In another example, the protrusions are locatedradially outward at between approximately 20% and 50% of the radiallength from a radial inner edge of the first and second web 106, 108.

The preceding description is exemplary rather than limiting in nature.Variations and modifications to the disclosed examples may becomeapparent to those skilled in the art that do not necessarily depart fromthe essence of this disclosure. The scope of legal protection given tothis disclosure can only be determined by studying the following claims.

What is claimed is:
 1. A turbine section for a gas turbine engine comprising: a first rotor assembly; a second rotor assembly located downstream of the first rotor assembly; a vane including a cooling passage having an outlet on a radial inner end of the vane; and a seal extending between the first rotor assembly and the second rotor assembly, the seal including a first plurality of openings in fluid communication with the cooling passage and a second plurality of opening directed at the second rotor assembly and spaced radially inward from the first plurality of openings, wherein the seal and the first rotor assembly at least partially define a first stage aft cavity and the seal and the second rotor assembly at least partially define a second stage forward cavity and the seal is configured to rotate with the first and second rotor assemblies.
 2. The turbine section of claim 1, wherein the seal is a segmented seal and includes a plurality of knife edges that mate with an abradable material on the radially inner end of the vane separating the first rotor assembly from the second rotor assembly.
 3. The turbine section of claim 2, wherein a radially inner edge of the abradable material is located between 10% and 50% of a radial length of a root portion of a blade in the first rotor assembly from a radial inner edge of the root portion.
 4. The turbine section of claim 1, wherein the first stage aft cavity includes a radially inner edge located between 10% and 50% of a radial length of a root portion of a blade in the first rotor assembly from a radial inner edge of the root portion.
 5. The turbine section of claim 4, wherein the second stage forward cavity includes a radially inner edge located between 10% and 50% of a radial length of a root portion of a blade in the second rotor assembly from a radial inner edge of the root portion of the blade in the second rotor assembly.
 6. The turbine section of claim 1, wherein the first rotor assembly includes a first web with a radially innermost protrusion located radially outward at between 30% and 60% of a radial length of the first web from a radial inner edge of the first web.
 7. The turbine section of claim 6, including a second stage having a second web with a radially innermost protrusion located radially outward at between 30% and 60% of a radial length of the second web from a radial inner edge of the second web.
 8. A gas turbine engine comprising: a turbine section including: a first rotor assembly; a second rotor assembly located downstream of the first rotor assembly; a vane including a cooling passage on a radial inner end; and a seal extending between the first rotor assembly and the second rotor assembly, the seal including a first plurality of openings in fluid communication with the cooling passage and a second plurality of opening directed at the second rotor assembly and spaced radially inward from the first plurality of openings, wherein the seal and the first rotor assembly at least partially define a first stage aft cavity and the seal and the second rotor assembly at least partially define a second stage forward cavity and the seal is configured to rotate with the first and second rotor assemblies.
 9. The gas turbine engine of claim 8, wherein the seal is a segmented seal and includes a plurality of knife edges that mate with an abradable material on the radially inner end of the vane separating the first rotor assembly from the second rotor assembly.
 10. The gas turbine engine of claim 9, wherein a radially inner edge of the abradable material is located between 10% and 50% of a radial length of a root portion of a blade of the first rotor assembly from a radial inner edge of the root portion.
 11. The gas turbine engine of claim 8, wherein the first stage aft cavity includes a radially inner edge located between 10% and 50% of a radial length of a root portion of a blade in the first rotor assembly from a radial inner edge of the root portion.
 12. The gas turbine engine of claim 11, wherein the second stage forward cavity includes a radially inner edge located between 10% and 50% of a radial length a root portion of a blade in the second rotor assembly from a radial inner edge of the root portion of the blade in the second rotor assembly.
 13. A method of operating a gas turbine engine comprising: directing bleed air from a first rotor assembly at a first pressure to purge a first stage aft cavity; and directing bleed air through a radially inner end of a vane at a second pressure to purge a second stage forward cavity and through a seal, wherein the first pressure is greater than the second pressure and the seal extends between a first rotor assembly and a second rotor assembly, the seal includes a first plurality of openings in fluid communication with a cooling passage in the vane and a second plurality of opening directed at the second rotor assembly and spaced radially inward from the first plurality of openings, wherein the seal and the first rotor assembly at least partially define the first stage aft cavity and the seal and the second rotor assembly at least partially define the second stage forward cavity and the seal is configured to rotate with the first and second rotor assembly.
 14. The method of claim 13, wherein a radially inner end of the vane includes a radially extending bleed air passage.
 15. The method of claim 13, wherein the first stage aft cavity includes a radially inner edge located between 10% and 50% of a radial length from a radial inner edge of a root portion of a blade in the first rotor assembly.
 16. The method of claim 15, wherein the second stage forward cavity includes a radially inner edge located between 10% and 50% of a radial length from a radial inner edge of a root portion of a blade in a second rotor assembly. 